GB2221979A - A combustion chamber for a gas turbine engine - Google Patents
A combustion chamber for a gas turbine engine Download PDFInfo
- Publication number
- GB2221979A GB2221979A GB8819537A GB8819537A GB2221979A GB 2221979 A GB2221979 A GB 2221979A GB 8819537 A GB8819537 A GB 8819537A GB 8819537 A GB8819537 A GB 8819537A GB 2221979 A GB2221979 A GB 2221979A
- Authority
- GB
- United Kingdom
- Prior art keywords
- apertures
- combustion chamber
- wall
- cooling
- rows
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/182—Transpiration cooling
- F01D5/184—Blade walls being made of perforated sheet laminae
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
1 A COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE The present invention
relates to combustion chambers for gas turbine engines, and is particularly concerned with cooling of the walls of the combustion chamber.
One conventional method of cooling the walls of combustion chambers of gas turbine engines uses cooling rings which are positioned between and secured to axially spaced wall sections. These cooling rings are provided with a plurality of relatively large apertures arranged in a row, or a number of rows of relatively small apertures. These apertures direct a flow of cooling fluid onto the inner surface of the wall to form a film of cooling fluid which protects the wall from the high temperatures produced in the combustion chamber. However, such cooling rings are relatively wasteful of cooling fluid.
A further problem with the cooling rings together tc is that the thermal gradients produced across the cooling ring lead to cracking of the cooling ring and the large numbers of cooling apertures allows easy propagation of the crack and eventual failure of the cooling ring.
A further conventional method of cooling the wall of combustion chambers of gas turbine engines uses walls which are formed from two or more laminae which are secured form internal passages therethrough for transpiration cooling of the wall by a cooling fluid. The cooling fluid is then directed through apertures out of the wall to from a cooling film of fluid on the inner surface of the wall. These arrangements are more efficient than the cooling rings using approximately a third of the cooling fluid, but the inner surface of the wall tends to become relatively hot because of ineffective film -cooling due to the apertures being arranged normal to the inner surface and being spaced by relatively large distances.
The present invention seeks to provide a combustion chamber of a gas turbine with improved film cooling of the walls of the combustion chamber.
1 2 Accordingly the present invention provides a combustion chamber for a gas turbine having at least one wall defining at least partially the combustion chamber, the wall having an inner surface and an outer surface, and additionally having at least one row of apertures extending therethrough for supplying cooling fluid onto the inner surface of the wall to form a cooling film of fluid on that surface, the axes of the apertures being arranged to form an angle of between 20 0 and 40 0 with the inner surface of the wall, each aperture having a first portion and a second portion, the first portion being arranged to receive cooling fluid from cooling fluid flowing over the outer surface of the wall and to supply the cooling fluid to the second portion, the second portion being divergent and arranged to direct the cooling fluid over the inner surface of the wall to form the cooling film of fluid.
The axes of the apertures may be arranged at an angle of between 25 0 and 35 0 with respect to the inner surface of the wall.
The divergent portions of the apertures may be divergent at an angle of substantially 12.5 0 with respect to the axes of the apertures.
The first portion of the apertures may by cylindrical.
The axes of the adjacent apertures in each row may be spaced apart by at least three times the diameter of the cylindrical portion of the apertures.
The wall may have at least two rows of apertures, the apertures in each row being staggered with respect to the apertures in the adjacent row or rows.
The adjacent rows of apertures may be spaced apart by at least two times the diameter of the cylindrical portion of the apertures.
The cylindrical portion of the apertures may have a diameter of substantially 0.762 MM.
The wall may be an upstream wall of the combustion chamber.
The wall maybe a tubular wall of a tubular combustion chamber, or may be an inner annular wall of an annular 3 combustion chamber, or may be annular combustion chamber.
An upstream portion of the wall may have the apertures arranged in axially spaced groups, each group having three rows of apertures.
A downstream portion of the wall may have the apertures arranged in axially spaced groups, each group having two rows of apertures.
The present invention will be more fully described by way of example with reference to the accompanying drawings, in which:
Figure 1 is partially cut away view of a gas turbine engine showing a combustion chamber according to the present invention.
Figure 2 is an enlarged longitudinal cross-sectional view of the combustion chamber shown in Figu're 1.
Figure 3 is an enlarged longitudinal cross-sectional view of an outer annular wall of the combustion chamber shown in Figure2.
Figure 4 is Figure 3.
Figure 5 is an enlarged longitudinal cross-sectional view of a portion of the outer annular wall shown in Figure 3.
Figure 6 is a view in the direction of arrow B in Figure 5.
Figure 7 is a view in the direction of arrow C in Figure 5.
Figure 8 is a cross-sectional view in the direction of arrows D-D, in Figure 5.
Figure 9 is an enlarged cross-sectional view through the upstream wall shown in Figure 2 in a - plane perpendicular to the plane of the sheet.
Figure 10 is a view in the direction of arrow E in Figure 9.
A turbofan gas turbine engine 10 is shown in Figure 1, and this comprises in axial flow series an inlet 12, a fan section 14, a compressor section 16, a combustor section an outer annular wall of a view in the direction of arrow A in 4 18, a turbine section 20 and an exhaust nozzle 22. The operation of the turbofan gas turbine engine 10 is quite conventional in that air flows into the inlet 12 and is given an initial compression by the fan section 14. This air is divided into two portions. The first portion of air is passed through the fan duct (not shown) to the f an nozzle (not shown). The second portion of air supplied to the compressor section 16 where the air is further compressed before being supplied to the combustor section 18. Fuel is burnt in the air supplied to the combustor section 18 to produce hot gases which f low through and drive the turbine section 20 before passing through the exhaust nozzle 22 to atmosphere. The turbine section 20 is arranged to drive the fan section 14 and compressor section 16 via shafts (not shown).
The combustor section 18 is shown more clearly in Figures 2 to 10. The combustor section comprises an outer casing 24 and an annular combustion chamber 26 enclosed by the casing 24. The annular combustion chamber 26 is defined by an annular upstream wall 28, an annular outer wall 30 and an -annular inner wall 32. An annular outer passage 25 for the flow of cooling air is formed between the casing 24 and the annular outer wall 30, and an inner passage 27 for the flow of cooling is formed within the annular inner wall 32.
The annular upstream wall 28 is provided with a plurality of equicircumferentially spaced apertures 36, and a fuel injector 34 is positioned coaxially in each of the apertures 36. The annular upstream wall 28 comprises an upstream wall member 37 and a downstream wall member 38 with a chamber 39 formed therebetween. The upstream wall member 37 has a plurality of apertures (not shown) for supplying air to the chamber 39. The downstream wall member 38 shown in Figure 9 and 10 is formed from a plurality of arcuate segments 54 each of which has a central aperture 40 formed substantially in its centre to receive a fuel injector 34. Each segment 54 is secured to the upstream wall member 37 by a number of bolts 64 and nuts (not shown).
an inner segments apertures air from J The segments 54 of the downstream wall member 38 have surface 56 and an outer surface 58, and the 54 are provided with a plurality of rows of 60 extending therethrough which supply cooling the chamber 39 onto the inner surface 56 of the segments 64 to form a cooling film of air. The rows of apertures 60 extend radially with respect to the axis of the annular combustion chamber 26. The apertures 60 are arranged so that their axes form an angle of between 20 0 and 40 0 with the inner surface 56 of the segments 54. The apertures 60, have first portions which are cylindrical, and second portions which are divergent. The cylindrical portions supply cooling air from the chamber 39 to the divergent portions, and the divergent portions direct the cooling air over the inner surface 56 of the segments 54 to form a cooling film of air. The divergent portions of the apertures diverge at an angle, in this example, of 12.5 0 with respect to the axes of the apertures. The axes of the adjacent apertures 60 in each row are spaced apart by three times the diameter of the cylindrical portion of the aperture.
It is to be noted that the rows of apertures 60 are arranged in groups of three rows, each group of rows of apertures being angularly spaced from the next group. The apertures in each row are staggered with respect to the apertures in the adjacent row or rows in that group.
The adjacent rows of apertures in each group are spaced apart by at least two times the diameter of the cylindrical portion of the apertures.
There are two groups of three rows of apertures 60 on one circumferential half of the segment 54, and another two groups of three rows of apertures 60 on the other circumferential half of the segment 54, these groups of apertures 60 are arranged to direct the cooling air in a circumferential direction towards the central aperture 40.
The outer annular wall 30 shown in figures 3 to 8 has an inner surface 44 and an outer surface 46, and has a plurality of rows of apertures 48. The apertures 48 extend 6 through the outer annular wall 30 to supply cooling air from the outer annular passage 25 onto the inner surface 44 of the outer annular wall 30 to form a cooling film of air. The rows of apertures 48 extend circumferentially with respect to the axis of the annular combustion chamber 26. The apertures 48 are arranged so that their axes from an angle of between 20 0 and 40 0 with respect to the inner surface of the outer annular wall 30. The apertures 48 have first portions 50 which are cylindrical, and second portions 52 which are divergent. The cylindrical portions 50 supply cooling air flowing over the outer surface 46 of the outer annular wall 30 in the outer annular passage 25 to the divergent portions 52, and the divergent portions 52 direct the cooling air in a downstream direction over the inner surface 44 of the outer annular wall 30 to form a cooling film of air. The divergent portions 52 of the apertures 48 diverge at an anglewp--12.5 0 with respect to the axes of the apertures 48. The axes of the adjacent apertures 48 in each row are spaced apart by a distance S, the distance S is three times the diameter d of the cylindrical portion 50 of the apertures 48. The divergent portions 52 of the apertures 48 diverge in a circumferential direction to produce a fan shaped aperture.
It is to be noted that the rows of apertures 48 are arranged in groups of three rows over an upstream portion 31 of the outer annular wall 30, and are arranged in groups of two rows over a downstream portion 33 of the outer annular wall 30. Each group of three rows of apertures in the upstream portion 31, or each group of two rows of apertures in the downstream portion 33 is axially spaced from the next group. The apertures 48 in each row are staggered with respect to the apertures 48 in the adjacent row or rows in that group.
The adjacent rows of apertures 48 in each group are spaced apart by at least two times the diameter d of the cylindrical portion 50 of the apertures 48.
Preferably the apertures 48 are arranged so that their axes form an angle of between 25 0 and 35 0 with respect to 1 the inner surface 44 of the outer annular wall 30.
The cylindrical portions 50 of the apertures 48 in this example have a diameter d of 0.762 mm, and the apertures are formed by laser drilling or other suitable method.
The spacing S, or Ditch, between the apertures is the most important dimension, and this is related to the angle of divergence of the apertures. The spacing S between the apertures increases with the angle of divergence of the apertures. In this example the angle oc of divergence of the apertures is 12.5 0 f and the spacing S is three times the diameter d. Apertures having angles c>< of greater than 12.5 0 will have a spacing S greater than three times the diameter d.
The apertures are inclined with respect to the inner surface of the upstream wall or annular outer wall so that the cooling air flowing through the apertures forms a cooling film of air on the inner surface of the upstream wall or annular outer wall. Apertures arranged at 90 0 to the inner surface of the walls do not form cooling films of air because the cooling air does not flow over the inner surface of the wall.
The apertures are divergent to improve the effectiveness of the cooling film of air by reducing the velocity of the air, causing the cooling air to spread out and merge with the cooling air from adjacent apertures in each row, and to ensure the cooling film remains on the inner surface of the walls.
However, with the single row of apertures although the effectiveness of cooling is improved, there is some entrainment of hot gases, produced in the combustion process, between the cooling film of air and the inner surface of the walls.
The use of several closely spaced rows of apertures arranged as a group is particularly beneficial, because the cooling film of air discharged over the inner surface of the wall by the first row of apertures acts as a barrier to inhibit the entrainment of hot gases between the cooling 8 film produced by the second row of apertures and the inner surface of the wall, and likewise the cooling films of air discharged over the inner surface of the wall by the second row of apertures acts as a further barrier to inhibit the entrainment of the hot gases between the cooling film produced by the third row of apertures and the inner surface of the wall. The use of several closely spaced rows of apertures produces a thicker cooling film of air which prevents the hot gases contacting the inner surface of the walls.
The use of walls with cooling apertures as described is more effective than the prior art cooling ring, because it uses a smaller amount of air to cool the same area, the invention uses approximately two thirds of the quantity of cooling air used by the prior art cooling ring.
The annular inner wall may also be provided with rows of apertures similarly arranged to the rows of apertures in the annular outer wall.
The invention although it has been described with reference to an annular combustion chamber may equally well be applied to tubular combustion chambers, or other arrangement of combustion chamber.
The rows of cooling apertures are simple to produce and they may be arranged at any location axial and/or circumferential to cope with local hot spots, ie local arrangements of rows of cooling apertures may positioned to provide film cooling for areas of the combustion chamber which are normally overheated.
The divergent portions of the adjacent apertures in each row are arranged such that the divergent portions do not merge together ie there is a space separating the divergent portions of the adjacent apertures in each. row.
9
Claims (16)
1. A combustion chamber for a gas turbine engine having at least one wall defining at least partially the combustion chamber, the wall having an inner surface and an outer surface, and additionally having at least one row of apertures extending therethrough for supplying cooling fluid onto the inner surface of the wall to form a cooling film of fluid on that surface, the axes of the apertures being arranged to form an angle of between 200 and 40 0 with the inner surface of the wall, each aperture having a first portion and a second portion, the first portion being arranged to receive cooling fluid from cooling fluid f lowing over the outer surface of the wall and to supply the cooling fluid to the second portion, the second portion being divergent and arranged to direct the cooling fluid over the inner surface of the wall to form the cooling film of fluid.
2. A combustion chamber as claimed in claim 1 in which the axes of the apertures are arranged at an angle of between 25 0 and 35 0 with respect to the inner surface of the wall.
3. A combustion chamber as claimed in claim 1 or claim 2 in which the divergent portions of the apertures are divergent at an angle of substantially 12.5 0 with respect to the axes of the apertures.
4. A combustion chamber as claimed in claim 1, claim 2 or claim 3 in which the first portions of the apertures are cylindrical.
5. A combustion chamber as claimed in claim 4 in which the axes of the adjacent apertures in each row are spaced apart by at least three times the diameter of the cylindrical portion of the apertures.
6. A combustion chamber as claimed in any of claims 1 to 5 in which the wall has at least two rows of apertures, the apertures in each row being staggered with respect to the apertures in the adjacent row or rows.
7. A combustion chamber as claimed in claim 6 when dependent upon claim 4 or claim 5 in which the adjacent rows of apertures are spaced apart by at least two times the diameter of the cylindrical portion of the apertures.
8. A combustion chamber as claimed in claim 4 in which the cylindrical portion of the apertures have a diameter of substantially 0.762 mm.
9. A combustion chamber as claimed in any of claims 1 to 8 in which the wall is an upstream wall of the combustion chamber.
10. A combustion chamber as claimed in any of claims 1 to 8 in which the wall is a tubular wall of a tubular combustion chamber.
11. A combustion chamber as claimed in any of claims 1 to 8 in which the wall is an inner annular wall of an annular combustion chamber.
12. A combustion chamber as claimed in any of claims 1 to 8 in which the wall is an outer annular wall of an annular combustion chamber.
13. A combustion chamber as claimed in any of claims 10 to 12 in which the wall has an upstream portion, the upstream portion having the apertures arranged in axially spaced groups, each group having three rows of apertures.
14. A combustion chamber as claimed in any of claims 1 to 13 in which the wall has a downstream portion, the downstream portion having the apertures arranged in axially spaced groups, each group having two rows of apertures.
15. A combustion chamber for a gas turbine engine substantially as hereinbefore described with reference to and as shown in Figures 2 to 10 of the accompanying.
16. A gas turbine engine comprising a combustion chamber as claimed in any of claims 1 to 15.
Published 1990 atTliePatentOffice. State House. 66171 High Holborn, London WC1R 4TP. Further copies maybe obtainedfrom, The PateritOffice Sales Branch, St Mary Gray. Orpingt0P. Kel.t BR5 3RD. Prired by Mikjplex techniques ittL St Mary Cray, Kent. Con. 1187
Priority Applications (6)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB8819537A GB2221979B (en) | 1988-08-17 | 1988-08-17 | A combustion chamber for a gas turbine engine |
| US07/374,837 US5000005A (en) | 1988-08-17 | 1989-07-03 | Combustion chamber for a gas turbine engine |
| JP1173862A JPH0275819A (en) | 1988-08-17 | 1989-07-05 | Combustion chamber for gas turbine engine |
| DE3924473A DE3924473A1 (en) | 1988-08-17 | 1989-07-24 | COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE |
| FR898910356A FR2635577B1 (en) | 1988-08-17 | 1989-08-01 | COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE |
| JP003664U JPH10279U (en) | 1988-08-17 | 1998-05-27 | Gas turbine engine combustion chamber |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB8819537A GB2221979B (en) | 1988-08-17 | 1988-08-17 | A combustion chamber for a gas turbine engine |
Publications (3)
| Publication Number | Publication Date |
|---|---|
| GB8819537D0 GB8819537D0 (en) | 1988-09-21 |
| GB2221979A true GB2221979A (en) | 1990-02-21 |
| GB2221979B GB2221979B (en) | 1992-03-25 |
Family
ID=10642266
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB8819537A Expired - Fee Related GB2221979B (en) | 1988-08-17 | 1988-08-17 | A combustion chamber for a gas turbine engine |
Country Status (5)
| Country | Link |
|---|---|
| US (1) | US5000005A (en) |
| JP (2) | JPH0275819A (en) |
| DE (1) | DE3924473A1 (en) |
| FR (1) | FR2635577B1 (en) |
| GB (1) | GB2221979B (en) |
Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5223320A (en) * | 1990-06-05 | 1993-06-29 | Rolls-Royce Plc | Perforated two layered sheet for use in film cooling |
| US5233828A (en) * | 1990-11-15 | 1993-08-10 | General Electric Company | Combustor liner with circumferentially angled film cooling holes |
| US5241827A (en) * | 1991-05-03 | 1993-09-07 | General Electric Company | Multi-hole film cooled combuster linear with differential cooling |
| US5261223A (en) * | 1992-10-07 | 1993-11-16 | General Electric Company | Multi-hole film cooled combustor liner with rectangular film restarting holes |
| US5279127A (en) * | 1990-12-21 | 1994-01-18 | General Electric Company | Multi-hole film cooled combustor liner with slotted film starter |
| EP0592161A1 (en) * | 1992-10-06 | 1994-04-13 | ROLLS-ROYCE plc | Gas turbine engine combustor |
| US5307637A (en) * | 1992-07-09 | 1994-05-03 | General Electric Company | Angled multi-hole film cooled single wall combustor dome plate |
| US5329761A (en) * | 1991-07-01 | 1994-07-19 | General Electric Company | Combustor dome assembly |
| US5465572A (en) * | 1991-03-11 | 1995-11-14 | General Electric Company | Multi-hole film cooled afterburner cumbustor liner |
| US5720434A (en) * | 1991-11-05 | 1998-02-24 | General Electric Company | Cooling apparatus for aircraft gas turbine engine exhaust nozzles |
| GB2441342A (en) * | 2006-09-01 | 2008-03-05 | Rolls Royce Plc | Wall Elements for Gas Turbine Engine Components |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS5312684U (en) * | 1976-07-15 | 1978-02-02 | ||
| GB9018014D0 (en) * | 1990-08-16 | 1990-10-03 | Rolls Royce Plc | Gas turbine engine combustor |
| GB9018013D0 (en) * | 1990-08-16 | 1990-10-03 | Rolls Royce Plc | Gas turbine engine combustor |
| CA2048726A1 (en) * | 1990-11-15 | 1992-05-16 | Phillip D. Napoli | Combustor liner with circumferentially angled film cooling holes |
| US5181379A (en) * | 1990-11-15 | 1993-01-26 | General Electric Company | Gas turbine engine multi-hole film cooled combustor liner and method of manufacture |
| GB9106085D0 (en) * | 1991-03-22 | 1991-05-08 | Rolls Royce Plc | Gas turbine engine combustor |
| US5435139A (en) * | 1991-03-22 | 1995-07-25 | Rolls-Royce Plc | Removable combustor liner for gas turbine engine combustor |
| US5220795A (en) * | 1991-04-16 | 1993-06-22 | General Electric Company | Method and apparatus for injecting dilution air |
| US5237813A (en) * | 1992-08-21 | 1993-08-24 | Allied-Signal Inc. | Annular combustor with outer transition liner cooling |
| DE4232442A1 (en) * | 1992-09-28 | 1994-03-31 | Asea Brown Boveri | Gas turbine combustion chamber |
| DE4427222A1 (en) * | 1994-08-01 | 1996-02-08 | Bmw Rolls Royce Gmbh | Heat shield for a gas turbine combustor |
| DE19502328A1 (en) * | 1995-01-26 | 1996-08-01 | Bmw Rolls Royce Gmbh | Heat shield for a gas turbine combustor |
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| US5758503A (en) * | 1995-05-03 | 1998-06-02 | United Technologies Corporation | Gas turbine combustor |
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| GB201518345D0 (en) * | 2015-10-16 | 2015-12-02 | Rolls Royce | Combustor for a gas turbine engine |
| DE102015225505A1 (en) | 2015-12-16 | 2017-06-22 | Rolls-Royce Deutschland Ltd & Co Kg | Wall of a component to be cooled by means of cooling air, in particular a gas turbine combustion chamber wall |
| US10823410B2 (en) * | 2016-10-26 | 2020-11-03 | Raytheon Technologies Corporation | Cast combustor liner panel radius for gas turbine engine combustor |
| US10670269B2 (en) * | 2016-10-26 | 2020-06-02 | Raytheon Technologies Corporation | Cast combustor liner panel gating feature for a gas turbine engine combustor |
| DE102017203326A1 (en) | 2017-03-01 | 2018-09-06 | Rolls-Royce Deutschland Ltd & Co Kg | Combustor shingle arrangement of a gas turbine |
| CN108194203A (en) * | 2017-12-19 | 2018-06-22 | 中国船舶重工集团公司第七0三研究所 | A kind of branch's cooling structure for industry gas turbine box-transfer story |
| US11867402B2 (en) * | 2021-03-19 | 2024-01-09 | Rtx Corporation | CMC stepped combustor liner |
| JP2024091028A (en) * | 2022-12-23 | 2024-07-04 | 川崎重工業株式会社 | Gas turbine combustor |
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| EP0187731A1 (en) * | 1985-01-07 | 1986-07-16 | United Technologies Corporation | Combustion liner for a gas turbine engine |
| EP0199534A1 (en) * | 1985-04-18 | 1986-10-29 | Ishikawajima-Harima Jukogyo Kabushiki Kaisha | Liner structure for a gas turbine combustion chamber |
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- 1989-07-05 JP JP1173862A patent/JPH0275819A/en active Pending
- 1989-07-24 DE DE3924473A patent/DE3924473A1/en not_active Withdrawn
- 1989-08-01 FR FR898910356A patent/FR2635577B1/en not_active Expired - Fee Related
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| EP0187731A1 (en) * | 1985-01-07 | 1986-07-16 | United Technologies Corporation | Combustion liner for a gas turbine engine |
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Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5223320A (en) * | 1990-06-05 | 1993-06-29 | Rolls-Royce Plc | Perforated two layered sheet for use in film cooling |
| US5233828A (en) * | 1990-11-15 | 1993-08-10 | General Electric Company | Combustor liner with circumferentially angled film cooling holes |
| US5279127A (en) * | 1990-12-21 | 1994-01-18 | General Electric Company | Multi-hole film cooled combustor liner with slotted film starter |
| US5465572A (en) * | 1991-03-11 | 1995-11-14 | General Electric Company | Multi-hole film cooled afterburner cumbustor liner |
| US5483794A (en) * | 1991-03-11 | 1996-01-16 | General Electric Company | Multi-hole film cooled afterburner combustor liner |
| US5241827A (en) * | 1991-05-03 | 1993-09-07 | General Electric Company | Multi-hole film cooled combuster linear with differential cooling |
| US5329761A (en) * | 1991-07-01 | 1994-07-19 | General Electric Company | Combustor dome assembly |
| US5720434A (en) * | 1991-11-05 | 1998-02-24 | General Electric Company | Cooling apparatus for aircraft gas turbine engine exhaust nozzles |
| US5307637A (en) * | 1992-07-09 | 1994-05-03 | General Electric Company | Angled multi-hole film cooled single wall combustor dome plate |
| EP0592161A1 (en) * | 1992-10-06 | 1994-04-13 | ROLLS-ROYCE plc | Gas turbine engine combustor |
| US5261223A (en) * | 1992-10-07 | 1993-11-16 | General Electric Company | Multi-hole film cooled combustor liner with rectangular film restarting holes |
| GB2441342A (en) * | 2006-09-01 | 2008-03-05 | Rolls Royce Plc | Wall Elements for Gas Turbine Engine Components |
| GB2441342B (en) * | 2006-09-01 | 2009-03-18 | Rolls Royce Plc | Wall elements with apertures for gas turbine engine components |
Also Published As
| Publication number | Publication date |
|---|---|
| FR2635577A1 (en) | 1990-02-23 |
| GB2221979B (en) | 1992-03-25 |
| JPH10279U (en) | 1998-12-04 |
| DE3924473A1 (en) | 1990-02-22 |
| FR2635577B1 (en) | 1994-06-17 |
| JPH0275819A (en) | 1990-03-15 |
| US5000005A (en) | 1991-03-19 |
| GB8819537D0 (en) | 1988-09-21 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 20000817 |